This application is based on and claims the priority under 35 U.S.C. xc2xa7119 of German Patent Application 100 05 348.3, filed on Feb. 8, 2000, the entire disclosure of which is incorporated herein by reference.
The invention relates to a method of fabricating a leading edge nose structure of an aerodynamic surface such as wing or a tail fin of an aircraft. The method involves bending or forming fuselage skin sheets and joining strengthening components onto the skin sheets.
Conventional wings of known aircraft types are fabricated as riveted and adhesively bonded aluminum constructions. The leading edge nose structure of the tail fin of such a known aircraft similarly is constructed of aluminum, glass fiber composites, graphite fiber composites, or carbon fiber composites, which are disadvantageously subject to erosion. In fabricating such known metal structures of aerodynamic surfaces, the fuselage skin sheets are formed over positive cores before the sheets are joined together. Other solutions or methods of fabricating the leading edge nose structure of an aerodynamic surface, and especially such a leading edge nose structure that allows boundary layer control by suction of boundary layer air through holes in the leading edge nose structure, have not been successfully developed or brought into practice. Nonetheless, the development of such improved methods of fabricating such leading edge nose structures is becoming ever more important in the field of aircraft construction, due to the constant efforts to reduce manufacturing costs and to improve the resulting aerodynamic structures with regard to an improvement of the aerodynamics, a reduction of the weight, and a resulting reduction of fuel consumption and increase of the maximum cruise flight range of the aircraft.
Previously known constructions for leading edge structures enabling a boundary layer suction all suffer from several common disadvantages, which has made it impossible or impractical to carry out such methods in large scale assembly line or series production of aircraft components. For example, the Boeing Company successfully tested a hybrid laminar technology in flight tests with the Boeing 757 aircraft. In that context, the connection of the several suction chambers was carried out by gluing or adhesively bonding together trapezoidal shaped metal sheets with relatively large joint surfaces on the aircraft outer skin.
Disadvantageously, the large joint surfaces of the adhesively bonded surfaces covered a rather large proportion of the suction holes, i.e. the aerodynamic surfaces provided with perforated holes therein. As a result, this known joining method has been very critically reviewed and judged, for example also with respect to the operating life and with respect to the application of internal pressure for cleaning and de-icing the suction structure.
Another known fabrication method involves warm deforming and joining of components in a vacuum soldering oven or furnace, using negative and positive cores made of steel. In this known method, considerable problems have arisen from the local and global deformations which require an extremely complicated and costly temperature control and variation. Additional difficulties have been caused by crystal transformations in the materials during the process. It has further been found in practice that the weight and the stiffness of the components cannot be brought down to acceptable values for making such a known method useful for the series production of a high volume of components.
Another known approach was the wind tunnel model of the so-called ELFIN II wing (an acronym for xe2x80x9cEuropean Laminar Flow Investigationxe2x80x9d), of which the components were joined together by adhesive bonding or gluing. While the results seemed acceptable for test purposes, such a fabrication method does not appear to be practical for high volume series production, because high manufacturing costs and a low operating life are expected. Also, on technical grounds the required interior pressurization for cleaning and for de-icing the suction structures is lacking.
In view of the above, it is an object of the invention to provide a method of fabricating a leading edge nose structure for an aerodynamic surface, especially for use in the manufacture of aircraft, which is able to form almost any desired aerodynamic contour with a high contour accuracy and true reproduction of the desired contour, while maintaining tight construction tolerances and achieving a high surface quality. The method particularly aims to avoid the use of cost-intensive and complex deformation techniques using positive mold cores or warm deformation of the material. To the contrary, the invention aims to reduce the manufacturing cost, effort and complexity, so as to enable the series production of a high volume of aerodynamic surface components. The invention further aims to avoid or overcome the disadvantages of the prior art, and to achieve additional advantages, as apparent from the present specification.
The above objects have been achieved according to the invention in a method of fabricating a leading edge nose structure of an aerodynamic surface, using a bending apparatus including two longitudinally movable guide elements movably arranged and supported on a threaded spindle, such that a rotation of the threaded spindle is transferred to a sliding motion of the two guide elements either toward each other or away from each other.
A respective forming tool member with a defined outer contour is articulately connected by a respective pivot joint to each respective one of the guide elements. The method according to the invention proceeds as follows.
Metal stringers or other profile element segments acting as strengthening members are arranged substantially parallel to each other and are then mechanically secured to a metal outer skin sheet that is in a flat planar configuration. For example, the strengthening members are laser welded or soldered onto the outer skin sheet so that the major profile web of each strengthening member stands substantially perpendicularly relative to the plane of the outer skin sheet. The bending apparatus is then configured and positioned so that the left and right forming tool members are arranged with their outer contour surfaces respectively contacting the back or inner contact surfaces of the strengthening members, while extending substantially crosswise across the strengthening members. At this point, the outer skin sheet is initially still in a substantially flat planar configuration.
Then respective outermost ones of the strengthening members are secured respectively to the left and right forming tool members by respective assembly pins. Next, the bending apparatus is operated to move the guide elements stepwise toward one another so that the left and right forming tool members pivot and slide relative to one another and thereby pull the outer skin sheet along with the moving forming tool members, so as to step-wise bend the outer skin sheet into a bent form that is generally in the shape of a parabola on a plane substantially perpendicular to the longitudinal extending strengthening members. This bent form of the outer skin sheet is determined by the similarly curved outer contour of the left and right forming tool members, against which the inner contact edge of each strengthening member is pulled, as well as a free-form parabolic curve of the outer skin sheet spanning the gap between the two forming tool members following the pivoting motion of the two forming tool members.
During this step-wise bending process, successive ones of the strengthening members are secured to the respective forming tool members by respective assembly pins, as each step of bending successively brings the next successive pair of strengthening members into contact with the forming tool members. This process is continued until the outer skin sheet has been bent or deformed into a curved aerodynamically contoured nose shape as seen in a cross-section plane perpendicular to the center line of the outer skin sheet or perpendicular to the strengthening members, in the area of the respectively adjacent free ends of the two forming tool members. The above described bending process is a cold bending or cold forming process, which may be carried out at room temperature (e.g. 60xc2x0 F. to 100xc2x0 F.).
The preliminary nose structure that has been joined together and then flexed or bent to apply a pre-tension thereto in the above described manner is then positioned into a negative form or mold, which has an inner contour substantially matching and following (while compensating tolerances) the generally parabolic curve contour of the outer skin sheet that has been formed as described above. Then, the two longitudinally extending edges of the outer skin sheet are secured to prevent the deformed nose structure from springing out of the negative form, by several securing plates which are distributed along the longitudinally extending edges of the negative form and which retain the edges of the outer skin sheet. Next, the above mentioned assembly pins are removed, to release the inner contact surfaces of the strengthening members from the left and right forming tool members, so that the bending apparatus can then be removed and retracted away from the nose structure.
Further steps according to the inventive method may then be carried out as follows, especially if the nose structure is to provide for boundary layer suction. After the preliminary nose structure has been arranged in the negative form or mold as described above, perforated throttling sheets, which have been pre-formed into a shape similar to a parabola generally matched or adapted to the parabolic contour of the outer skin sheets, are then arranged in a row and butted against one another in the longitudinal direction on the inside of the preliminary nose structure, i.e. contacting the inner contact edges of the strengthening members, after having applied a sealant at defined locations between the inner perforated throttling sheets and the strengthening members, wherever required. The inner perforated throttling sheets are then secured to the inner contact edges of the strengthening members by respective blind rivets. The non-abutting crosswise extending edge of the last perforated throttling sheet in the row of these sheets ends and extends perpendicularly above or inwardly along the skin sheet crosswise edge, while the longitudinal edges of the throttling sheets extend perpendicularly inwardly along the longitudinally extending edges of the skin sheets.
Next, a plurality of inner ribs that have been previously bent into a parabolic shape matching that of the throttling sheets, are provided with a sealant on the back or outwardly directed surfaces thereof and are then placed onto the inwardly facing surface of the throttling sheets so as to extend generally crosswise relative to the longitudinal centerline of the leading edge nose structure. These inner ribs are then secured by respective blind rivet connections.
Then, a plurality of flanged U-section members having a xe2x80x9chat-shapedxe2x80x9d sectional profile are provided with a sealant applied thereto, and are arranged running substantially parallel to one another along the longitudinal direction of the leading edge nose structure, on the inner parabolic surface of the inner throttling sheets, and are secured thereto by additional blind rivet connections. Thereafter, a plurality of longitudinally extending inner skin sheets are provided with a sealant along their longitudinally extending edges, and then these edge portions are arranged on the protruding flanges of the flanged U-section members and connected thereto by additional blind rivet connections. In the above manner, the leading edge nose sandwich structure has been essentially completed with a sandwich construction, which is then removed out of the negative form or mold after releasing the securing plates.